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The Mars Scout Phoenix Lander successfully landed in the Martian northern latitude on May 25, 2008. The Robotic Arm, which was designed to dig and to transfer soil samples to other Lander instruments, contained a number of actuators that had specific operational windows on the Martian surface due to the bearing lubricant. The deployment of the Robotic Arm was planned for Sol 2 (Mars days are referred to “Sols”). A few weeks before Mars landing, the Robotic Arm operations team learned that a strict flight rule had been imposed. It specified that the deployment shall be accomplished when the actuators were at or above −25°C since the deployment activity was qualified with the actuators at −40°C. Furthermore, the deployment plan identified a window of opportunity between 13:00 Local Solar Time (LST, equivalent to dividing the Sol into 24 equal Martian hours) and 15:30 LST.

The Third Generation ENose is an air quality monitor designed to operate in the environment of the US Lab on the International Space Station (ISS). It detects a selected group of analytes at target concentrations in the ppm regime at an environmental temperature range of 18 – 30 °C, relative humidity from 25 – 75% and pressure from 530 to 760 torr. This device was installed and activated on ISS on Dec. 9, 2008 and has been operating continuously since activation. Data are downlinked and analyzed weekly. Results of analysis of ENose monitoring data show the short term presence of low concentration of alcohols, octafluoropropane and formaldehyde as well as frequent short term unknown events.

The Mars Science Laboratory will be the next Martian mobility system that is scheduled to launch in the fall of 2011. The ChemCam Instrument is a part of the MSL science payload suite. It is innovative for planetary exploration in using a technique referred to as laser breakdown spectroscopy to determine the chemical composition of samples from distances of up to about 9 meters away. ChemCam is led by a team at the Los Alamos National Laboratory and the Centre d'Etude Spatiale des Rayonnements in Toulouse, France. The portion of ChemCam that is located inside the Rover, the ChemCam Body Unit contains the imaging charged-coupled device (CCD) detectors. Late in the design cycle, the ChemCam team explored alternate thermal design architectures to provide CCD operational overlap with the Rover's remote sensing instruments. This operational synergy is necessary to enable planning for subsequent laser firings and geological context.

The Mars Science Laboratory (MSL) mission to land a large rover on Mars is being prepared for Launch in 2011. A Multi-Mission Radioisotope Thermoelectric Generator (MMRTG) on the rover provides an electrical power of 110 W for use in the rover and the science payload. Unlike the solar arrays, MMRTG provides a constant electrical power during both day and night for all seasons (year around) and latitudes. The MMRTG dissipates about 2000 W of waste heat to produce the desired electrical power. One of the challenges for MSL Rover is the thermal management of the large amount of MMRTG waste heat. During operations on the surface of Mars this heat can be harnessed to maintain the rover and the science payload within their allowable limits during nights and winters without the use of electrical survival heaters. A mechanically pumped fluid loop heat rejection and recovery system (HRS) is used to pick up some of this waste heat and supply it to the rover and payload.

NASA's proposed lunar lander, Altair, will be exposed to vastly different external temperatures following launch till its final destination on the moon. In addition, the heat rejection is lowest at the lowest environmental temperatures (0.5 kW @ 4K) and highest at the highest environmental temperature (4.5 kW @ 215K). This places a severe demand on the radiator design to handle these extreme turn-down requirements. A radiator with digital turn-down capability is currently under study at JPL as a robust means to meet the heat rejection demands and provide freeze protection while minimizing mass and power consumption. Turndown is achieved by independent control of flow branches with isolating latch valves and a gear pump to evacuate the isolated branches. A bench-top test was conducted to characterize the digital radiator concept. Testing focused on the demonstration of proper valve sequencing to achieve turn-down and recharge of flow legs.

Launched on India's Chandrayaan-1 spacecraft on October 22, 2008, JPL's Moon Mineralogy Mapper (M3) instrument has successfully completed over six months of operation in space. M3 is one in a suite of eleven instruments, six of which are foreign payloads, flying onboard the Indian spacecraft. Chandrayaan-1, managed by the Indian Space Research Organization (ISRO) in Bangalore, is India's first deep space mission. Chandrayaan-1 was launched on the upgraded version of India's Polar Satellite Launch Vehicle (PSLV-XL) from the Satish Dhawan Space Centre, SHAR, Sriharikota, India. The primary science objective of the M3 instrument is the characterization and mapping of the lunar surface composition in the context of its geologic evolution. Its primary exploration goal is to assess and map the Moon mineral resources at high spatial resolution to support future targeted missions.

Previous studies indicated evidence of opportunistic pathogens in samples obtained during missions to the International Space Station (ISS). This study utilized TaqMan quantitative PCR to determine specific gene abundance in potable and non-potable ISS waters. Probe and primer sets specific to the small subunit rRNA genes were designed and used to elucidate overall bacterial rRNA gene numbers. In addition, primer-probe sets specific for Burkholderia cepacia and Stenotrophomonas maltophilia were optimized and genes of these two opportunistic pathogens quantified in the pre- and post-flight drinking water as well as coolant waters. This Q-PCR approach supports findings of previous culture-based studies however; the culture based studies may have underestimated the microbial burden of ISS drinking water.

Launched on NASA's Aura spacecraft on July 15, 2004, JPL's Tropospheric Emission Spectrometer (TES) has been operating successfully for over three years in space. TES is an infrared high resolution, imaging fourier transform spectrometer with spectral coverage of 3.3 to 15.4 μm to measure and profile essentially all infrared-active molecules present in the Earth's lower atmosphere. It measures the three-dimensional distribution of ozone and its precursors in the lower atmosphere on a global scale. The Aura spacecraft was successfully placed in a sun-synchronous near-circular polar orbit with a mean altitude of 705 km and 98.9 minute orbit period. The observatory is designed for a nominal 5 year mission lifetime. The instrument thermal design features include four temperature zones needed for efficient cryogenic staging to provide cooling at 65 K, 180 K, 230 K and 300 K.

The Moon Mineralogy Mapper (M3) instrument is scheduled for launch in 2008 onboard the Indian Chandrayaan-1 spacecraft. The mission is managed by the Indian Space Research Organization (ISRO) in Bangalore, India and is India's first flight to the Moon. M3 is being developed for NASA by the Jet Propulsion Laboratory under the Discovery Program Office managed by Marshall Space Flight Center. M3 is a state-of-the-art instrument designed to fulfill science and exploratory objectives. Its primary science objective is to characterize and map the lunar surface composition to better understand its geologic evolution. M3's primary exploration goal is to assess and map the Moon mineral resources at high spatial resolution to support future targeted missions. M3 is a cryogenic near infrared imaging spectrometer with spectral coverage of 0.4 to 3.0 μm at 10 nm resolution with high signal to noise ratio, spatial and spectral uniformity.

The Orbiting Carbon Observatory (OCO) instrument is scheduled for launch onboard an Orbital Sciences Corporation LEOStar-2 architecture spacecraft in December 2008. The instrument will collect data to identify CO2 sources and sinks and quantify their seasonal variability. OCO observations will permit the collection of spatially resolved, high resolution spectroscopic observations of CO2 and O2 absorption in reflected sunlight over both continents and oceans. OCO has three bore-sighted, high resolution, grating spectrometers which share a common telescope with similar optics and electronics. A 0.765 μm channel will be used for O2 observations, while the weak and strong CO2 bands will be observed with 1.61 μm and 2.06 μm channels, respectively. The OCO spacecraft circular polar orbit will be sun-synchronous with an inclination of 98.2 degrees, mean altitude of 705 km and 98.9 minute orbit period.

The advent of lightweight fairings for new spacecraft and the increased thrust of new launch vehicles have intensified the need for better techniques for predicting and for reducing the low frequency noise environment of spacecraft at lift-off. This paper presents a VAPEPS (VibroAcoustic Payload Environment Prediction System) parametrical analysis of the noise reduction of spacecraft fairings and explores a novel technique for increasing the low frequency noise reduction of lightweight fairings by approximately 10 dB.

A Small Loop Heat Pipe (SLHP) featuring a wick of only 1.27 cm (0.5 inches) in diameter has been designed for use in spacecraft thermal control. It has several features to accommodate a wide range of environmental conditions in both operating and non-operating states. These include flexible transport lines to facilitate hardware integration, a radiator capable of sustaining over 100 freeze-thaw cycles using ammonia as a working fluid and a structural integrity to sustain acceleration loads up to 30 g. The small LHP has a maximum heat transport capacity of 120 Watts with thermal conductance ranging from 17 to 21 W/°C. The design incorporates heaters on the compensation chamber to modulate the heat transport from full-on to full-stop conditions. A set of start up heaters are attached to the evaporator body using a specially designed fin to assist the LHP in starting up when it is connected to a large thermal mass.

Radiators will be used on Mars to reject excess heat from various processes and surfaces and will help temper the climate of any future manned habitats. Radiator performance is a function of the radiator size (area), the emissivity, ε, of the radiator surface, the radiator temperature, local environmental conditions, and the effective sky temperature to which it radiates. The effective sky temperature of Mars is not known. Previous estimates have ranged between 80 K to 170 K. Also, it is not known how dust accumulation and other environmental effects act to change the performance of a radiator as a function of time. The MTERC Experiment is designed to gather data to address these unknowns. This paper will describe the operational theory and the configuration of the MTERC experiment hardware and will discuss results of MTERC performance testing.

The Multi-angle Imaging SpectroRadiometer (MISR) instrument was launched aboard NASA’s Earth Observing System (EOS) Terra spacecraft on December 18, 1999. The overall mission design lifetime for the instrument is 6 years. The EOS Terra spacecraft was placed in a sun-synchronous near-circular polar orbit with an inclination of 98.3 degrees and a mean altitude of 705 km. The overall objective of MISR is to provide a means to study the ecology and climate of Earth through the acquisition of global multiangle imagery on the daylit side of Earth. MISR views the sunlit Earth simultaneously at nine widely spaced angles, collects global images with high spatial detail in four colors at every angle. The images acquired, once calibrated, provide accurate measurements of brightness, contrast and color of reflected sunlight.

This paper describes a study on the capillary limit of a loop heat pipe (LHP) with two evaporators and two condensers. Both theoretical analysis and experimental investigation are performed. Experimental tests conducted include heat load to one evaporator only, even heat loads to both evaporators, and uneven heat loads to both evaporators. Test results show that after the capillary limit is exceeded, vapor will penetrate through the wick of the weaker evaporator, and the compensation chamber (CC) of that evaporator will control the loop operating temperature regardless of which CC has been in control prior to the event. Because the evaporator can tolerate vapor bubbles, the loop can continue to work after vapor penetration. As the loop operating temperature increases, the system pressure drop actually decreases due to a decrease in liquid and vapor viscosities. Thus, the loop may reach a new steady state at a higher operating temperature after vapor penetration.

One of the new technologies successfully demonstrated on the recent Mars Pathfinder mission was the active Heat Rejection System (HRS). This system consisted of a mechanically pumped cooling loop, which actively controlled the temperatures of the various parts of the spacecraft. A single phase Refrigerant 11 liquid was mechanically circulated through the lander and cruise electronics box heat exchangers. This liquid transferred the excess heat to an external radiator on the cruise stage. This is the first time in unmanned spacecraft history that an active heat rejection system of this type has been used on a long duration spacecraft mission. Pathfinder was launched in December 1996 and landed on the Martian surface on July 4, 1997. The system functioned flawlessly during the entire seven months of flight from Earth to Mars. A life test set up of the cooling loop was used to verify the life of the system.

In October, 1998, the first of the NASA New Millennium Spacecraft, DS1, was successfully launched into space. The objectives for this spacecraft are to test advanced technologies that can reduce the cost or risk of future missions. One of these technologies is the Solar Concentrator Array with Refractive Linear Element Technology (SCARLET). Although part of the advanced technology validation study, the array is also the spacecraft power source. Funded by BMDO, the SCARLET™ concentrator solar array is the first spaceflight application of a refractive lens concentrator. As part of the DS1 validation process, the amount of array diagnostics is very extensive. The data obtained includes temperature measurements at numerous locations on the 2-wing solar array. For each individual panel, a 5-cell module in one of the circuit strings is wired so that a complete I-V curve can be obtained. This data is used to verify sun pointing accuracy and array output performance.

This paper reports several slow reversible and quasi-reversible processes which occur in the porous electrode/solid electrolyte combination at AMTEC operating temperatures. These processes help to elucidate the evolution of the electrode and electrolyte characteristics with time. They also demonstrate that the atomic constituents of the electrode/electrolyte engage in significant dynamic motion. We report the stability of the sodium beta“-alumina phase in low pressure sodium vapor at 1173K up to 3000 hours, and the decomposition of the sodium meta-aluminate (NaAlO2) phase present at about 1% in the BASE ceramic, which gives rise to transient local increases in the solid electrolyte resistivity due to local micro-cracking. We also report slow apparent morphological changes, possibly surface or grain boundary reconstruction, in TiN and RhW electrodes driven by changes in the local sodium activity.

A miniature quadrupole mass spectrometer array and gas chromatograph have been designed and built for NASA flight missions. Without the gas chromatograph the mass spectrometer is to be used for detection, by astronauts in EVA, of N2, O2, the hydrazines, and NH3 leaks in the hull of the International Space Station, and of adsorbed hydrazines on the astronauts’ suits. The fully-adapted astronaut system, with all software and visual readout, is called the Trace Gas Analyzer. When interfaced with the miniature gas chromatographic system, the mass spectrometer will be useful for a variety of NASA missions involving more complex gas mixtures. The missions include planetary exploration (to Venus, Europa, Titan, etc.), as well as cabin-air monitoring for long-duration human flight to the Moon, Mars, and beyond.

Mars Exploration Rover (MER) mission launched two spacecraft to Mars in June and July of 2003 and landed two rovers on Mars in January 2004. A Heat Rejection System (HRS) based on a mechanically pumped single-phase liquid cooling system was used to reject heat from electronics to space during the seven months cruise from Earth to Mars. Even though most of this HRS design was similar to the system used on Mars Pathfinder in 1996, several key modifications were made in the MER HRS design. These included the heat exchanger used in removing the heat from electronics, design of venting system used to vent the liquid prior to Mars entry, inclusion of pressure transducer in the HRS, and the spacecraft radiator design. Extensive thermal/fluids modeling and analysis were performed on the MER HRS design to verify the performance and reliability of the system. The HRS design and performance was verified during the spacecraft system thermal vacuum tests.